Gas turbine engines are widely used in commercial aircraft throughout the world. These engines include a forward stage of fan blades that are rotatably mounted to the front of a turbine engine and are driven at high speeds by the engine. The turbine engine is generally surrounded by an aerodynamically shaped engine shroud that provides a protective cover for the engine and also increases aerodynamic efficiency. In order to further increase damage tolerance and aerodynamic efficiency, the fan blades, engine and engine shroud may be surrounded by a ducted nacelle. The inner diameter of the nacelle is sized to encompass and form a duct around the fan blades and engine shroud. The nacelle forces the high velocity air stream produced by the fan blades to flow between the inner surface of the nacelle and the outer surface of the engine shroud. The flow of air produced by the fan blades is generally referred to as the fan or bypass air flow and contributes to the propulsive force produced by the engine.
The compression and turbine stages of the turbine engine also produce a high velocity air flow that contributes to overall engine propulsion. The high velocity air flow produced by the turbine engine flows out the aft end of the engine rearward of the nacelle. Unlike the bypass air flow, however, the flow produced by the turbine engine is at a mix of air and combusted fuel byproducts at a highly elevated temperature.
During an engine failure, such as a bleed duct failure, explosion, or overpressurization failure, it is possible for the high temperature flow of gases within the turbine engine to vent into the space between the engine and the engine shroud or other compartment in the installation. During a bleed duct or other failure, a significant volume of high-temperature air flow into the space between the engine and engine shroud or other installation. This flow of gases pressurizes and heats the space between the turbine engine and engine shroud. If left to its own course, the high temperatures and pressures caused by a bleed duct or other failure can cause a structural failure of the engine shroud or other installation.
Federal aircraft regulations, FAR 25.1103, requires that engine installations safely withstand the effects of an engine bleed duct failure. In order to comply with FAR 25.1103 and prevent failure of the engine structure during a bleed duct or other failure, it is common to locate a series of pressure relief doors around the exterior of the engine shroud aft of the nacelle, on the inlet of the nacelle and on the engine pylon. The pressure relief or "blowout doors" are rectangularly shaped to fit rectangularly shaped contents. The pressure relief doors are mechanically mounted to the shroud or other structure so that they automatically open in case of excessive compartment pressurization. During a bleed duct or other failure, the pressure relief doors open and allow the high temperature gases to flow out of the shroud thus preventing overpressurization and subsequent structural failure.
Although the pressure relief doors prevent structural failures due to overpressurization, the structure downstream of the doors must still be able to withstand contact with the high temperature gases escaping through the pressure relief doors. After hot engine gases escape through the pressure relief doors, they flow rearward along the exterior surface of the structure. The hot engine gases can remain in contact with the exterior surface of the structure for a sufficient amount of time to elevate the temperature of the surrounding structure. Under some conditions, it is possible for the surrounding structure to be elevated to a sufficiently high temperature to cause structural failure due to overheating.
To prevent such structural failures, the structure aft of the pressure relief doors has traditionally been protected by thermal insulating blankets. Alternatively, the structure has been fabricated of a material capable of withstanding the high temperatures produced by the escaping engine gases, for example, titanium. The use of either thermal insulating blankets or high temperature materials present a number of disadvantages. Using insulating blankets complicates the fabrication and maintenance of the engine installation and adds to the overall cost and weight. The use of titanium and other high temperature materials can also add weight to the overstructure and also add to fabrication complexity, difficulty, and cost. In addition, coefficient of expansion differences between a titanium or steel structure and the surrounding aluminum or composite structure can add to the complexity of the overall installation.
As seen from the above discussion, there exists a need in the aircraft industry for a pressure relief system to safely address the problems created by an engine bleed duct or other failure while reducing the disadvantages of prior art solutions. The present invention is directed to such a system.